Silicon carbide (SiC)-based ceramic matrix composite (CMC) materials have been proposed as materials for certain components of gas turbine engines, such as the turbine blades and vanes. Various methods are known for fabricating SiC-based CMC components, including Silicomp, melt infiltration (MI), chemical vapor infiltration (CVI), polymer inflation pyrolysis (PIP), and oxide/oxide processes. Though these fabrication techniques significantly differ from each other, each involves the use of hand lay-up and tooling or dies to produce a near-net-shape part through a process that includes the application of heat at various processing stages.
As with turbine blades and vanes formed from more conventional superalloy materials, CMC blades and vanes are primarily equipped with cavities and cooling passages to reduce weight, reduce centrifugal load, and reduce operating temperatures of the components. These features are typically formed in CMC components using a combination of removable and expendable tooling.
Current fabrication methods for forming internal passages or cavities in CMC components use materials that need to be “melted” out or removed from the internal passage during the burn-out cycle. Forming CMC component with a cavity includes a number of steps, including using pre-forms. First, a plurality of ceramic plies, some of which can include reinforcing material or are pre-impregnated with matrix, are laid up on a mandrel or mold in a pre-determined pattern to provide desired final or near-net-shape and desired mechanical properties of component. The mandrel is generally selected from materials such as tin, vinyl, or other meltable materials. The laid-up plies may be pre-impregnated (pre-preg) with matrix material, such as SiC or impregnated with matrix after lay-up of the plies. Prior to densification of the CMC pre-form, the mandrel is removed through a burn-out cycle. In the burn-out cycle, the CMC pre-form is inverted and the mandrel forming materials, such as tin, vinyl, or other meltable materials are melted out through the open tip of the pre-form CMC leaving an open tip area.
After the burn-out cycle, the CMC pre-form blade is very fragile due to burn-off of the volatile substances of the composite. The open tip area of the CMC pre-form requires capping or closing before use in gas turbines. In known processes, to close the open tip area of the CMC pre-form a tip cap is inserted into the fragile open tip area. The tip cap can be formed form of a CMC laminate part having a number of plies, generally 20-50 plies, and shaped as the open tip area to fill the open tip area of the CMC pre-form. Forming the CMC laminate tip cap by cutting out the CMC plies to the desired shape and laying up the plies in the desired geometry is time and labor intensive. Challenges also arise with placing the CMC laminate having a number of plies into the open tip area. Additionally, because both the CMC laminate and pre-form blade are fragile prior to densifying, these components can be easily damaged during assembly.
Therefore, methods of forming ceramic matrix composite components and ceramic matrix composite components for gas turbines that do not suffer from the above drawbacks is desirable in the art.